Conventionally, a rotorcraft commonly includes a main lift rotor provided with a plurality of blades.
The blades of the main lift rotor describe a very flat cone having a plane of rotation that is perpendicular to the general lift generated by said main rotor. This general lift of the main rotor may then be resolved into a vertical lift force proper and a horizontal force that serves to cause the helicopter to move in translation. Consequently, the main rotor does indeed provide the rotorcraft with lift.
In addition, on a helicopter, by controlling the shape and the tilt of said cone relative to the frame of reference of the helicopter, a pilot can control the advance of the helicopter and direct it accurately.
In order to act on the cone, the blades are caused to flap so as to modify their angle of inclination relative to the drive plane of the main rotor, said drive plane being perpendicular to the mast of the main rotor. By varying the pitch of a blade, the lift it generates is varied, thereby causing the flapping of said blade to vary.
Consequently, the helicopter is provided with specific means for causing the pitch of each blade to vary, and consequently for varying the aerodynamic angle of attack of each blade relative to the incident stream of air through which the blade passes.
In order to control the general lift of the main rotor, both in magnitude and in direction, the helicopter pilot thus acts generally on the value of the pitch angle of each blade by causing the blade to turn about its longitudinal pitch axis. Thus, when the pilot orders a collective variation of pitch, i.e. an identical variation in pitch for all of the blades, the pilot causes the magnitude of the lift generated by the main rotor to vary, thereby controlling the altitude and the speed of the helicopter.
In contrast, collective pitch variation has no effect on the direction of this general lift.
In order to modify the direction of the general lift generated by the main rotor, it is necessary to incline said cone by varying pitch in a manner that is not collective, but rather that is cyclic. Under such circumstances, the pitch of a blade varies as a function of its azimuth angle and during one complete revolution it passes between a maximum value to a minimum value that are obtained at opposite azimuth angles.
Cyclic variation of the pitch of the blades gives rise to cyclic variation in the lift from the blades and thus to variation in the angle of inclination of the cone. By cyclically varying the pitch of the blades, the pilot controls the attitude of the aircraft and its movement in translation.
The pilot's pitch flight controls, a collective pitch lever and a cyclic stick, are generally connected to the blades via mechanical connections known as “linkages”, which linkages are secured to the non-rotating plate of a cyclic swashplate. The rotating plate of the cyclic swashplate is mechanically connected to each blade via a respective pitch control rod.
More precisely, primary roll and pitch linkages connect the cyclic stick to a mixer, the mixer being connected to the non-rotary plate of the cyclic swashplate by secondary linkages. In addition, the collective pitch lever is connected to the mixer via a collective primary linkage. Under such circumstances, a flight control is connected to a rotor via a roll and pitch linkage or collective linkage, both of which are provided in succession with a primary linkage and a secondary linkage.
A movement of the cyclic stick causes the primary roll or pitch linkage to move and consequently causes the corresponding secondary linkage(s) to move via the mixer.
In contrast, a movement of the collective pitch lever gives rise to a movement of the collective primary linkage and then of the secondary linkage via the mixer.
Nevertheless, since the forces that need to be exerted are large, a servo-control is generally made available in each secondary linkage. For example, in a light helicopter, there is provided in principle one servo-control for pitch control, referred to as the “pitch servo-control” for convenience, and two servo-controls for roll to left and to right for piloting in roll.
When the pilot seeks to modify the collective pitch of the blades, action on the collective lever causes all three servo-controls to raise or to lower the cyclic swashplate as a whole, i.e. both the non-rotary plate and the rotary plate of the cyclic swashplate.
The pitch control rods are then all moved through the same distance, which implies that the pitch of all of the blades varies through the same angle.
In contrast, in order to apply cyclic pitch variation to the blades so as to direct the helicopter in a given direction, the pilot causes at least one servo-control to move by tilting the cyclic stick appropriately in the desired direction.
The cyclic swashplate then does not move vertically, but instead tilts relative to the mast of the main rotor. Each pitch control rod is thus moved as a function of the intended target so as to generate appropriate cyclic variation of the pitch of each blade.
Furthermore, the linkages are provided with at least one link rod and at least one crank means for connecting the pilot's flight controls to the servo-controls. In addition, provision may also be made for a phasing unit that enables the cyclic swashplate to tilt about two mutually perpendicular axes for use in heavy helicopters.
Similarly, the flight controls of a helicopter include a yaw control connected to a tail rotor by a yaw linkage that passes via the mixer.
The linkages are thus generally very long and heavy, particularly the yaw control linkage.
The members of the linkages give rise to friction forces that may be considerable for linkages of great length. Under such conditions the pilot may have difficulty in moving the cyclic stick or the collective pitch lever, given the amount of force required.
A first solution consists in using electric flight controls as suggested in document WO 2005/002963 or US 2007/0102588 (now U.S. Pat. No. 7,229,046). Nevertheless, that first solution is difficult to implement, in particular on existing rotorcraft.
Consequently, helicopter manufacturers have remedied the problem as posed by adding assistance systems that may be hydraulic or pneumatic. Known assistance systems consist in a block of actuators acting merely as a force-multiplying relay unit, with the block of actuators being arranged for example between bottom crank means and the mixer, for example.
Nevertheless, such assistance systems are bulky and heavy. Furthermore, they run the risk of hydraulic or pneumatic leakage, leading to a loss of effectiveness.
Finally, the gas in pneumatic assistance systems is sensitive to variations in temperature, unfortunately all too frequent in aviation, in particular because of operating at different altitudes, and the fluids used in hydraulic assistance systems possess polluting chemical compounds.
In addition, rotorcraft conventionally include an autopilot system for stabilizing the rotorcraft and/or for reducing the pilot's workload.
Fitting an autopilot system may possibly lead to “series” actuators being put into place in series in each linkage between a flight control and the controlled rotor.
The series actuators are intended for stabilizing the rotorcraft. Thus, the series actuators are generally very fast, but with little authority in amplitude.
Document EP 1 037 130 presents a “series” actuator suitable for being controlled by a computer to stabilize the machine.
A sensor sends information about an order given by a control stick to the computer. The computer then makes use of piloting relationships to control the series actuator in order to stabilize the helicopter.
Furthermore, an actuator known as a “parallel” actuator, or indeed as a “trim” actuator, is placed in parallel with each linkage.
Unlike series actuators, parallel actuators are controlled by autopilot means to control the movements of the helicopter. Thus, parallel actuators are generally slow, but have considerable authority in amplitude.
The autopilot controls the parallel actuators to maneuver the rotorcraft.
Furthermore, parallel actuators are used as anchors for series actuators. The autopilot continuously activates the series actuators in order to stabilize the rotorcraft. Without some additional action, the movements of the series actuators would be fed back to the pilot's flight controls.
By anchoring the associated flight controls, the parallel actuators prevent the movements of the series actuator from being fed back to the flight controls, e.g. the cyclic stick.
The parallel actuators also serve to anchor the flight controls.
In a first variant, a parallel actuator includes a motor suitable for moving an outlet shaft that is connected in parallel to a linkage. Between the motor and the outlet shaft, the parallel actuator is provided in particular with gearing, a spring box generating a determined force relationship, and a safety pin.
In another variant described in patent FR 2 708 112, the parallel actuator includes in particular in succession: a motor, a reversible gearing, a position sensor, a safety device, and an outlet shaft.
Control means activated by an autopilot then control the motor to generate a force relationship.
According to document US 2005/0173595 (now U.S. Pat. No. 7,108,232), the parallel actuator is an actuator suitable for generating a force relationship by being controlled by dedicated means.
Furthermore, dampers are provided in the linkages in order to stiffen the flight controls as a function of the speed at which the flight controls are moved.
Consequently, the state of the art includes rotorcraft having a piloting assistance device, provided in particular with a hydraulic block or with a plurality of series and parallel actuators.
Such a device is necessarily bulky, heavy, and expensive.